Axial flow compressors



April 13, 1965 F. w. w. MORLEY ETAL 3,178,101

AXIAL FLOW COMPRESSORS 2 Sheets-Sheet 1 Filed April 27, 1962 A tforneysApril 1965 F. w. w. MORLEY ETAL 3,178,101

AXIAL FLOW COMPRESSORS Filed Apri 1962 2 Sheets-Sheet 2 wag M ATTORNEYSUnited States Patent This invention relates to axial flow compressorblading particularly, but not exclusively, for compressors of gasturbine jet reaction engines.

While the invention may be applied either to stator or rotor blading,more important applications may be to rotor blading.

According to one aspect of the present invention there is provided anaxial flow compressor blade having a working portion of aerofoilcross-section, a root portion at one end of said working portion, and astrengthening insert incorporated in said root portion saidstrengthening insert lying in a spanwise sense and being curved so thatit lies substantially parallel to the mean profile camber radius of theworking portion.

The strengthening insert may be formed from a metal or a thermosettingresinous material and the remaining portions of the blade may be formedfrom reinforced thermosetting resinous material. Thus the strengtheninginsert may be formed from magnesium or aluminium.

In a preferred arrangement the reinforced thermosetting resinousmaterial comprises fibres of glass or asbestos impregnated with aphenolic or an epoxy resin. Preferably the reinforced thermo-settingresinous material extends the full length of the blade from the tip ofthe blade to the root portion and encloses the strengthening insert.

The strengthening insert is preferably substantially triangular incross-section, the lower wall of the triangle being contained in theplane of the lower wall of the root portion.

The root portion of the blade is preferably dovetail shape incross-section.

According to the present invention in another aspect an axial flowcompressor rotor is provided with compressor blades formed fromreinforced thermosetting resinous material, each blade incorporating inits root portion a metal strengthening insert, the strengthening insertsof some of the compressor blades being formed from a heavier materialthan the inserts of the remaining compressor blades so that, by suitablydisposing the blades with heavier strengthening inserts around theperiphery of the rotor, the assembly of rotor and compressor blades canbe dynamically balanced.

Blades according to the present invention may form part of an axial-flowcompressor of a gas turbine engine adapted to be mounted with itslongitudinal axis vertical so as to produce an upward component ofthrust on the aircraft in which it is mounted.

Experience in reinforced synthetic resin rotor blades has shown thatsuch blades are susceptible to failure should the reinforcing fibreshave to be severely bent during moulding to achieve any given shape,particularly at the intersection of a cambered blade profile with itsplatform or root.

By curving the strengthening insert in the manner set forth above, thebending or kinking" is minimised, while the angle through which radialfibres must be bent to fit over the strengthening insert is kept to aminimum.

One embodiment of the present invention will now be ice described by wayof example, with reference to the accompanying drawings in which- FIGURE1 is a side elevation of a gas turbine engine with parts broken away toshow the compressor rotor thereof,

' FIGURE 2 is an axial-section through part of the said compressorrotor, the said compressor rotor including rotor blades incorporatingthe present invention,

FIGURE 3 is a view in the direction of arrow 3 indicated on FIGURE 2,

FIGURE 4 is a section through the rotor blade taken on the line 44indicated in FIGURE 2,

FIGURE 5 is a section taken on the line 5-5 indicated on FIGURE 4,

FIGURE 6 is a section taken on the line 6-6 indicated on FIGURE 4,

FIGURE 7 is a section taken on the line 77 indicated on FIGURE 4, and

FIGURE 8 is a section taken on the line 8-8 indicated on FIGURE 2.

The gas turbine engine shown in FIGURE 1 comprises an axial flowcompressor 10 which receives air from an air-intake 11 and deliverscompressed air to combustion equipment 12 where fuel is burned in theair, the products of combustion passing through a turbine section 13before passing into a jet-pipe 14 and being exhausted to atmospherethrough a propelling nozzle 15.

The turbine section 13 is arranged to drive a compressor rotor 16forming part of the compressor 10 through shafting (not shown). Thecompressor rotor 16 includes axially-spaced rows of compressor rotorblades 17 which alternate with axially-spaced rows of stator blades 18secured to the casing of the compressor 10.

Each compressor rotor blade 17 is formed from reinforced thermo-settingresinous material such as glass or asbestos fibres impregnated with aphenolic or an epoxy resin and comprises a working portion 17a ofaerofoil cross-section and a root portion 17b by which the blade issecured to the compressor rotor disc 16.

The root portion 17b is of dovetail shape in crosssection and it isreceived in a correspondingly shaped slot 19 formed in the periphery ofthe rotor disc 16. Each rotor blade 17 is retained axially in positionon the rotor disc 16 by means of synthetic resin lug 20 and a metal tab21. The lug 20, which is provided at one end of the root portion 17b,contacts the upstream face of the rotor disc 16. The metal tab 21 liesbetween the lower end of the root portion 17b and the lower wall of theslot 19, one end of the tab 21 having an outwardly directed flange 21awhich engages the lug 20, and the other end of the tab 21 having aradially-inwardly directed flange 21b which contacts the face of therotor disc 16.

Each root portion 17b is strengthened by means of a strengthening insert22 which is disposed within the root portion 17b in a spanwise sense andwhich is curved in a plane transverse to the radial plane of the workingportion The radius of curvature of each strengthening insert 22 is thesame as the mean profile camber radius of the working portion 17a. Themean profile camber radius is shown by the chain dotted line 23indicated on FIGURE 4.

The strengthening inserts 22 are substantially triangular incross-section, the lower wall of the triangle being contained in theplane of the lower wall of the respective root portion 17b.

As will be seen more clearly in FIGURES 5 to 7 the glass fibres 24 ofthe blade material extend the full length of the blade 17 from the tipthereof to the root portion and enclose the strengthening insert 22. Theremainder of each root portion 17b is built up with pads 25 of syntheticresin material to give the straight dovetail shape.

resinous material. 7

In order to balance the compressor rotor disc 16 dynamically, some ofthe rotor blades 17 of a blade row maybe provided with magnesium inserts.22 and the remaining rotor blades of the blade.row.ma.y be providedwith aluminium inserts 22 which are suitably disposed around theperiphery of the rotor'disc '16.

We claim:

1. In an axialflow compressor rotorhaving a plurality of radiallyoutwardly extending blades, said blades having Working portions ofaerofoil cross-section,- and-having root portions at one end of saidworking portions, said workingportions having curved mean profilecamberradii and being formed from reinforced thermo-setting-resinous material,each of said blades having a metallic strengthening insert imbeddedwithin andsecured to ithe root portion of said blade, saidmetallicstrengthening insert lying in a chordwise sense and being curved so thatit lies parallel to the meanprofile camber radius of the workingportion, said metallic strengthening inserts of the plurality ofbiadesbeing formed from varying density metals with the inserts-of some of theblades being of heavier metal than the inserts of the remaining blades,whereby-the as Q, sembly of rotor and compressor blades can hedynamically balanced by suitably disposing the blades with the heavierstrengthening inserts around the periphery ofthe rotor.

2.1m an axial flow compressor rotor as claimed in claim 1, said heaviermetal inserts comprising aluminum and the remaining metal insertscomprising magnesium.

ReferencesCited by the Examiner UNITED STATES PATENTS 1,035,364 8/12Leblanc 25377 1,172,947 2/16 Coppage 74-573 2,346,552 4/44 Bro'tz -1592,454,200 11/48 Perkins 1701'59 2,775,426 12/56 Barretttet al. 230-134 2,830,647 4/58 Warnken 230-134 2,868,441 1/59 Reutt 230134 2,929,755 3/60jPorter -253--77 3,132,841 5/64 Wilder 253-77 FOREIGN PATENTS 7 124,1555/47 Australia.

233,906 5/61 Australia.

861,978 11/40 France. 1,232,125 4/60 France.

778,667 7 7/57 Great Britain.

KARL J. ALBRECHT, Primary Exazhirier.

'JCISEPH H. BRANSON, JR., Examiner.

1. IN AN AXIAL FLOW COMPRESSOR ROTOR HAVING A PLURALITY OF RADIALLYOURWARDLY EXTENDING BLADES, SAID BLADES HAVING WORKING PORTIONS OFAEROFOIL CROSS-SECTION, AND HAVING ROOT PORTIONS AT ONE END OF SAIDWORKING PORTIONS, SAID WORKING PORTIONS HAVING CURVED MEAN PROFILECAMBER RADII AND BEING FORMED FROM REINFORCED THERMO-SETTING RESINOUSMATERIAL, EACH OF SAID BLADES HAVING A METALLIC STRENGHTENING INSERTIMBEDDED WITHIN AND SECURED TO THE ROOT PORTION OF SAID BLADE, SAIDMETALLIC STRENGTHENING INSERT LYING IN A CHORDWISE SENSE AND BEINGCURVED SO THAT IT LIES PARALLEL TO THE MEAN PROFILE CAMBER RADIUS OF THEWORKING